This invention relates generally to gas turbine engines, and, more specifically, to a new and improved gas turbine engine having improved efficiency.
A significant goal for advancing gas turbine engine technology is providing improved engine thermal efficiency. One measure of efficiency is the engine's energy output divided by fuel energy input, which may be represented, for example, by specific fuel consumption (SFC), which is the ratio of fuel flow in pounds per hour divided by thrust of the engine in pounds.
The overall efficiency of an engine is affected by the efficiency of its various components. One significant engine component which substantially affects efficiency is the turbine. A conventional turbine includes one or more alternating rows of stationary stator nozzle blades and rotating turbine blades and may also include one or more turbine rotors, for example a high pressure turbine (HPT) driving a compressor in serial flow relationship with a low pressure turbine (LPT) driving, for example either a fan or low pressure compressor.
Modern day and advanced gas turbine engines operate at relatively high combustion gas temperatures for reducing SFC. Such relatively high temperatures typically require cooling of turbine blades, which is accomplished by bleeding a portion of compressor air and channeling it through the turbine blades for cooling. Inasmuch as this cooling air is bypassed from the main flowpath of the engine for cooling, a reduction in overall engine efficiency necessarily occurs.
A gas turbine engine is typically designed for obtaining a desired amount of work from the turbine thereof. Relatively high efficiency and relatively low quantities of cooling air as described above are two of many conventional goals utilized in designing the turbine.
Other goals utilized in designing turbines include relatively high performance and thrust; relatively low weight, cost and fuel flow; simplicity; and small size. Although meeting all these goals is desirable, actual design practice requires tradeoffs between them.
Also utilized in designing turbines are many conventional given turbine specifications including, for example, fluid temperatures and pressures at blade inlets and outlets, turbine power output required and shaft speeds. Velocity vector diagrams of fluid flow through turbine blade rows are then conventionally chosen at a preferred radius, such as at either the blade, hub (i.e. 0 percent blade height) or blade pitch/meanline radius (i.e. 50 percent blade height). The velocity diagrams typically include the velocity vectors of fluid flow at turbine blade inlets and outlets.
Velocity vector diagrams at other radial locations of the blades are next conventionally determined consistent with, among other things, radial equilibrium of the fluid flowing thereover. Radial equilibrium is a condition wherein the radial pressure force on the fluid is equal and opposite to the centrifugal force acting on the fluid due to tangential components of velocity therein.
The shape and size of the blades, including angular orientation of all sections thereof, are next conventionally generated from the velocity vector diagrams for defining the entire blade outer surface. Of course, additional conventional practices are also utilized for finally defining a preferred turbine design.
Reaction is a conventionally known parameter useful in defining the type of a turbine. Reaction has many alternative definitions including, for example, the percent stage static enthalpy drop occurring in a turbine rotor, and may be expressed in terms of temperature, Pressure or velocity parameters. Inasmuch as reaction may be expressed in terms of velocity, it follows that reaction may also be used as an indication of velocity vector diagrams and, therefore, as an indication of blade shape and orientation.
Two fundamental and conventionally known types of turbine blading include reaction blades and impulse blades. All gas turbines have blades which vary in the degree of reaction from hub to tip as a result of radial equilibrium conditions as indicated above. Inasmuch as reaction necessarily increases from hub to tip, a single reaction value, such as that at the pitch or meanline, is typically utilized to define the type of turbine.
A pure impulse turbine (i.e. 0 percent reaction) has blades which are generally symmetrical, crescent-shaped airfoils having a generally uniform channel between adjacent blades for obtaining equal inlet and outlet areas and fluid velocities. A reaction turbine has blades which are unsymmetrical, having relatively thick leading edge portions and thin trailing edge portions, adjacent ones of which define a converging channel for accelerating a fluid therebetween for obtaining a higher outlet velocity than the inlet velocity. In an impulse turbine no static pressure drop exists across the blades thereof, and in a reaction turbine a static pressure drop from inlet to outlet is effected.
Conventional turbines have pitch reactions ranging from about 10 percent up to about 50 percent. Reactions of 40 percent to 50 percent usually result in optimum performance or peak efficiency for a turbine stage according to two prior art references. One of those references also teaches peak efficiency when velocity vector diagrams are symmetrical.
Although the prior art teaches that optimum performance may be obtained at 40-50 percent reaction, relatively high reaction includes some negative tradeoffs as well. For example, increasing reaction increases the discharge swirl angle of gases leaving a turbine which must be accommodated by increasing the turning ability of downstream blades. Not only does increased swirl angle result in a more complex downstream blading, it also increases aerodynamic losses of the highly-turned gas flow.
Increased reaction also increases the acceleration, discharge Mach number and pressure drop of gases being channeled through the turbine blades. Inasmuch as aerodynamic efficiency losses are proportional to the square of velocity, relatively high reactions may result in relatively high mixing losses of blade trailing edge aperture discharge air. Also, the increased pressure drop will cause increased leakage of gas flow over blade tips.
Conventional turbines typically also include a compressor discharge, or thrust balance, seal for reducing differential internal thrust forces to a level that is manageable with conventional thrust bearings. More specifically, air is discharged from the last compressor rotor of a gas turbine engine at a first pressure which acts over an outlet area of the compressor, resulting in a forward generated force. Combustion gases at the inlet of the turbine rotor section of the engine are at a second pressure and act across an inlet area of the turbine for generating an aft-directed force. The forward force is substantially larger than the aft force, which is one reason for requiring the use of a thrust bearing to accommodate differential thrust acting on the compressor-turbine shaft. The compressor discharge seal is typically provided between the compressor and the combustor to reduce the area over which the compressor discharge pressure would otherwise act for reducing the forward thrust force. Inasmuch as the compressor discharge seal adds weight and complexity to an engine, it would be desirable if its use were not required.